A gas turbine engine includes a turbomachinery core having a high pressure compressor, combustor, and high pressure turbine (“HPT”) in serial flow relationship. The core is operable in a known manner to generate a primary gas flow. The high pressure turbine includes annular arrays (“rows”) of stationary vanes or nozzles that direct the gases exiting the combustor into rotating blades or buckets. Collectively one row of nozzles and one row of blades make up a “stage”. Typically two or more stages are used in serial flow relationship. These components operate in an extremely high temperature environment, and must be cooled by air flow to ensure adequate service life.
HPT nozzles are often configured as an array of airfoil-shaped vanes extending between annular inner and outer bands which define the primary flowpath through the nozzle.
Due to operating temperatures within the gas turbine engine, it may be desirable to utilize materials with high temperature capability and additionally low coefficient of thermal expansion. For example, to operate effectively in such strenuous temperature and pressure conditions, composite materials have been suggested and, in particular for example, ceramic matrix composite (CMC) materials. These low coefficient of thermal expansion materials have higher temperature capability than metallic parts. The higher operating temperatures within the engine result in higher engine efficiency and these materials may be lighter weight than traditionally used metals. However, such ceramic matrix composite (CMC) have mechanical properties that must be considered during the design and application of the CMC. CMC materials have relatively low tensile ductility or low strain to failure when compared to metallic materials. Also, CMC materials have a coefficient of thermal expansion which differs significantly from metal alloys used as restraining supports or hangers for CMC type materials.
A fairing assembly is located between the high pressure turbine and low pressure turbine within a gas turbine engine. The fairing forms a flow path between these high pressure and low pressure turbines and is generally made of metallic castings. However, it may be desirable to utilize a fairing assembly of which at least a portion is formed of a low coeffiecient of material which may be desirable due to its light weight, low coefficient of thermal expansion and high temperature capability.
Turbine center frame fairings are generally made of nickel based cast metallic alloys and use of a ceramic material may provide desirable weight savings which result in improved efficiency and performance of the gas turbine engine.
However, problems associated with dissimilar materials create relative growth problems between the adjacent connected parts. It may be desirable to overcome these and other deficiencies known in fairing assemblies in order to reduce fuel consumption, as well as potentially increasing operating temperature capability for gas turbine engines.
The information included in this Background section of the specification, including any references cited herein and any description or discussion thereof, is included for technical reference purposes only and is not to be regarded subject matter by which the scope of the disclosure is to be bound.